In conventional solar powered spacecraft, power is generated by a solar photovoltaic array (SP) and flows into a power management and distribution (PMAD) system. From there it is distributed to all loads: the bus loads, the payload(s) and in electric propulsion (EP) spacecraft, the EP load such as Hall effect thrusters. A PMAD system is a large, heavy, complex collection of circuits including e.g. filters, batteries, DC/DC converters, isolation circuits and voltage regulators. The EP load may use the largest share of the power when it is operating. This means that the PMAD system must have the capacity to process all the power produced by the SP including power for the EP even though significant amount of the time the PMAD system is servicing only the bus load(s) and payload(s). Thus the PMAD system has to be quite large to supply the required power to the EP on demand and also has the have the means to dissipate (as heat) excess power generated by the SP when the EP is making no demand. The extra size and weight required to perform both these functions is critical in spacecraft design and operation.
There is yet another shortcoming associated with current PMAD systems. The PMAD system may contain peak power tracking or solar array shunt circuitry in order to optimize the power provided to operate the EP and other loads. At the beginning of life (BOL) the SP provides greater peak power than at the end of life (EOL). The PMAD system peak power tracking unit (operating in continuous mode or discrete steps) lowers the voltage as the SP ages. However, the EP must always remain at a power level that is below the temporary maximum power, never at the maximum power, in order to preserve an operational stability margin to account for the unknown and unpredictable aging of the SP. Assume, for illustration sake that there are no other loads on the SP: the EP operating power point dictates the output power of the array which must be below the peak power point to provide for the operational stability margin. This means that the EP thruster does not get all the power it could get and hence the thrust is reduced and spacecraft transit/maneuver time is extended. In one prior attempt to improve solar powered spacecraft power system the power processor unit conditions the power output from the solar array and regulates the bus voltage and current so as to provide an output current applicable to be used on an arcjet thruster. U.S. Pat. No. 5,604,430. In another approach a power control circuit employs a multiplier, differentiator, detector, phase comparator and integrator to effect a ramp generator to produce minor variations in a beam current reference signal to an ion thruster. U.S. Pat. No. 4,143,314.